Dual flow turboshaft engine and improved hot flow nozzle

ABSTRACT

According to the invention, the hot flow nozzle includes a single skin ( 14   s ) conformed for promoting the cold flow ( 10 ) circulation when the aircraft including said turboshaft engine in a cruise flight mode.

The present invention relates to improvements to aircraft turbojetengines.

More specifically, the present invention relates to bypass (dual flow)turbojet engines comprising:

-   -   a nacelle delimiting, at the front, an air intake and axially        containing a cold flow fan and a hot flow generator provided        with a turbine enclosed in a casing ending, at the rear, in a        bulbous boss connected to said turbine casing along a first        connecting surface;    -   an engine cowl containing said hot flow generator and        comprising:        -   an external wall delimiting with said nacelle a cold flow            duct ending, at the rear, in a cold flow nozzle, and        -   an internal wall delimiting with said turbine casing a hot            flow duct,    -   a hot flow nozzle extending said engine cowl rearward by being        connected thereto along a second connecting surface and        collaborating with said boss to delimit a nominal nozzle throat        section and a nominal outlet section for said hot flow, and    -   at least one ventilation opening positioned between said        external wall of said engine cowl and said hot flow nozzle, on        the outside with respect to said cold flow nozzle, and intended        to discharge to the outside a ventilation air flow bled from        said cold flow and introduced into said engine cowl in order to        regulate the temperature of said hot flow generator.

In turbojet engines of this type, said hot flow nozzle is thereforesubjected to the cold flow on the external side and to the hot flow onthe internal side, and comprises an external skin that is optimized forthe flow of the cold flow and an internal skin that is optimized for theflow of the hot flow.

A double-skinned structure such as this is therefore relatively heavyand, in addition, because of the substantial temperature differencebetween the hot flow and the cold flow, behaves like a bimetallic stripand undergoes deformations that detract from the optimization of theflows of the hot and cold flows in flight, particularly in cruisingflight.

It is an object of the present invention to overcome thesedisadvantages.

To this end, according to the invention, a turbojet engine of the typerecalled hereinabove is noteworthy in that:

-   -   said hot flow nozzle consists of a single skin;    -   said single skin is shaped to encourage said cold flow to flow        when the aircraft is in cruising flight; and    -   the shape of said bulbous boss is tailored in order to be able        to respect said nominal nozzle throat section and said nominal        outlet section for said hot flow.

The present invention is based on the following observations:

-   -   that for most of the time that an aircraft turbojet engine is in        use it is in cruising flight and that it is therefore        advantageous to optimize the rear part of such a turbojet engine        for cruising flight;    -   that in cruising flight, the flow of the cold flow of such a        turbojet engine is a supersonic flow, whereas the flow of the        hot flow is a subsonic flow; and    -   that optimizing the flow of the supersonic cold flow is more        critical than optimizing the flow of the subsonic hot flow, and        that optimizing the flow of said cold flow can be taken almost        exclusively into consideration provided that the nominal nozzle        throat section and nominal outlet section are respected for said        hot flow.

Thus, according to the present invention, these observations are put togood use to create the hot flow nozzle in the form of a single skin,something which, by comparison with a double-skin embodiment, provides aweight saving and avoids bimetallic strip effect deformations.

In order to avoid excessive disruption to the hot flow, it is alsoadvantageous to ensure that the shape of said single skin avoids anyseparation of said hot flow.

Advantageously, along the axis of the turbojet engine, the shape of saidsingle skin varies continuously. Said single skin is bulbous in shape,for example, widening from said second connecting surface into thecontinuation of said external wall of said engine cowl.

Of course, the position of said ventilation opening, between the rearedge of the external wall of the engine cowl and said part of saidsingle skin continuing this external wall, depends on the shape of saidsingle skin.

As a preference, said ventilation opening is created in the vicinity ofthe largest-diameter part of said single skin.

Obviously, the present invention is particularly easy to implement inturbojet engines in which said hot flow nozzle throat coincides withsaid hot flow outlet, because then only one nominal section has to berespected.

The figures of the attached drawing make it easy to understand how theinvention may be embodied. In these figures, identical references denotesimilar elements.

FIG. 1 is a schematic view with cutaway, partially in axial section, ofa turbojet engine affected by the present invention.

FIG. 2 shows, in an enlarged and partial axial section, a knownconfiguration for the rear part of the hot flow duct of the turbojetengine of FIG. 1.

FIG. 3 shows, in a view comparable with that of FIG. 2, one example of aconfiguration according to the present invention, for the rear part ofsaid hot flow duct of the turbojet engine of FIG. 1.

In FIG. 2, said rear part according to the present invention has beensuperposed, in chain line, on said known configuration for comparisonpurposes.

Symmetrically, and likewise for the purposes of comparison, FIG. 3depicts said known rear part in chain line superposed on saidconfiguration according to the present invention.

The bypass turbojet engine 1 of axis L-L, depicted schematically andpartially in FIG. 1 and intended to power an aircraft (not depicted),comprises a nacelle 2 delimiting, at the front, an air intake 3. Thenacelle axially contains a cold flow fan 4 and a hot flow generator 5.The hot flow generator 5 is provided with a turbine enclosed in a casing6 ending, at the rear, in a bulbous boss 7, and fixed to said casing 6along a connecting surface or joining plane P1. The hot flow generator 5is enclosed in a cowl 8 comprising an external wall 8E and an internalwall 81 (see FIG. 2).

The external wall BE of the cowl 8 delimits with the nacelle 2 a duct 9for the cold flow, symbolized by the arrows 10, said cold flow duct 9ending in a cold flow nozzle 11. The internal wall 8I delimits, with theturbine casing 6, a duct 12 for the hot flow, symbolized by the arrows13.

The cowl 8 of the hot flow generator 5 is extended rearward by a hotflow nozzle 14 that collaborates with said boss 7 to extend said duct 12as far as the annular outlet orifice 15 for the hot flow 13. The hotflow nozzle 14 is attached to the rear end of the internal wall 81 alonga connecting surface or joining plane P2 and created between the rearedge 8R of the external wall 8E and said nozzle 14 is an annular opening16 positioned on the outside with respect to the cold flow nozzle 11 andintended to discharge to the outside a flow of ventilation airsymbolized by the arrows 17, bled from the cold flow 10 and introducedinto the cowl 8 of the hot flow generator 5 in order to regulate thetemperature thereof.

In the known embodiment depicted in solid line in FIG. 2, it is assumedthat the annular outlet orifice 15 for the hot flow 13 additionallyforms the throat of the nozzle 14. Furthermore, the latter consists ofan external wall 14E over which the cold flow 10 flows and of aninternal wall 141 over which the hot flow 13 flows.

When the aircraft bearing the turbojet engine 1 is in cruising flight,the cold flow 10 is supersonic, whereas the hot flow 13 is subsonic.

The object of the present invention illustrated schematically by FIG. 3is chiefly to lighten the hot flow nozzle 14 and prevent thedeformations thereof that are due to the bimetallic strip effect of saidwalls 14E and 141 while at the same time not in any way detracting fromthe performance of said turbojet engine when the aircraft is in cruisingflight.

To do that, said nozzle 14 consists of a single skin 14 s shaped toencourage the flow of the supersonic cold flow 10 when the aircraft isin cruising flight. The shape of the single skin 14 s variescontinuously along the axis L-L and can be likened to a bulbous shapewidening from the joining plane P2 into the continuation of the externalwall BE of the engine cowl 8, creating therewith an opening 16 scomparable to the opening 16.

In order, as far as the hot flow is concerned, to respect the nominalsection of the nozzle throat and the nominal outlet section—whichsections, in the example depicted, coincide with the hot flow outletorifice 15 s—the shape of the boss 7 is modified as depicted as 7 s.Thus, inside the nozzle 14, the hot flow duct 12 adopts the shape 12 s.

The shape of the single skin 14 s has to be such that, in its concaveportion facing toward the boss, there is no separation of said hot flow13.

It may be that, in order to give the single skin 14 s its optimal shape,the ventilation opening 16 has to be shifted to 16 s, as depicted inFIG. 3. The position 16 s of said ventilation opening is advantageouslyin the vicinity of the largest-diameter part of the single skin 14 s.

1-9. (canceled)
 10. A bypass turbojet engine (1) for an aircraft,provided with a longitudinal axis (L-L) and comprising: a nacelle (2)delimiting, at the front, an air intake (3) and axially containing acold flow (10) fan (4) and a not flow (13) generator (5) provided with aturbine enclosed in a casing (6) ending, at the rear, in a bulbous boss(7) connected to said turbine casing (6) along a first connectingsurface (P1); an engine cowl (8) containing said hot flow generator (5)and comprising: an external wall (8E) delimiting with said nacelle (2) acold flow duct (9) ending, at the rear, in a cold flow nozzle (11), andan internal wall (81) delimiting with said turbine casing (6) a hot flowduct (12), a hot flow nozzle (14) extending said engine cowl (8)rearward by being connected thereto along a second connecting surface(P2) and collaborating with said boss (7) to delimit a nominal nozzlethroat section and a nominal outlet section for said hot flow (13), andat least one ventilation opening (1E) positioned between said externalwall (8E) of said engine cowl (8) and said hot flow nozzle (14), on theoutside with respect to said cold flow nozzle (11), and intended todischarge to the outside a ventilation air flow (17) bled from said coldflow (10) and introduced into said engine cowl (8) in order to regulatethe temperature of said hot flow generator (5), wherein: said hot flownozzle (14) consists of a single skin (14 s); said single skin (14 s) isshaped to encourage said cold flow (10) to flow when the aircraft is incruising flight; and the shape of said boss (7) is tailored to suit thatof said single skin in order to be able to respect said nominal nozzlethroat section and said nominal outlet section for said hot flow (13).11. The turbojet engine as claimed in claim 10, wherein the shape ofsaid single skin (14 s) prevents separation of said hot flow in itsconcave portion.
 12. The turbojet engine as claimed in claim 10,wherein, along the axis (L-L) of said turbojet engine (1), the shape ofsaid single skin (14 s) varies continuously.
 13. The turbojet engine asclaimed in claim 12, wherein said single skin (14 s) is at leastapproximately bulbous in shape, widening from said second connectingsurface (P2) into the continuation of said external wall (8E) of saidengine cowl (8).
 14. The turbojet engine as claimed in claim 10, whereinthe position of said ventilation opening (16 s) is dependent upon theshape of said single skin (14 s).
 15. The turbojet engine as claimed inclaim 13, wherein said ventilation opening (16 s) is created in thevicinity of the largest-diameter part of said single skin (14 s). 16.The turbojet engine as claimed in claim 10, wherein said hot flow nozzlethroat coincides with said hot flow outlet (15 s).
 17. A hot flow nozzle(14) for a bypass turbojet engine (1) for an aircraft, provided with alongitudinal axis (L-L) and comprising: a nacelle (2) delimiting, at thefront, an air intake (3) and axially containing a cold flow (10) fan (4)and a hot flow (13) generator (5) provided with a turbine enclosed in acasing (6) ending, at the rear, in a bulbous boss (7) connected to saidturbine casing (6) along a first connecting surface (P1); an engine cowl(8) containing said hot flow generator (5) and comprising: an externalwall (8E) delimiting with said nacelle (2) a cold flow duct (9) ending,at the rear, in a cold flow nozzle (11), and an internal wall (81)delimiting with said turbine casing (6) a hot flow duct (12), said hotflow nozzle (14) extending said engine cowl (8) rearward by beingconnected thereto along a second connecting surface (P2) andcollaborating with said boss (7) to delimit a nominal nozzle throatsection and a nominal outlet section for said hot flow (13), and atleast one ventilation opening (16) positioned between said external wall(8E) of said engine cowl (8) and said hot flow nozzle (14), on theoutside with respect to said cold flow nozzle (11), and intended todischarge to the outside a ventilation air flow (17) bled from said coldflow (10) and introduced into said engine cowl (8) in order to regulatethe temperature of said hot flow generator (5), wherein: said hot flownozzle (14) consists of a single skin (14 s) shaped to encourage saidcold flow (10) to flow when the aircraft is in cruising flight.
 18. Thehot flow nozzle (14) as claimed in claim 17, wherein said single skin(14 s) is at least approximately bulbous in shape, widening from saidsecond connecting surface (P2) into the continuation of said externalwall (8E) of said engine cowl (8).